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At least 145 records · Page 8
All sourcesNASA NTRSDOE OSTI
NASA NTRS1980

Development of a vortex-lift-design method and application to a slender maneuver-wing configuration

A method has been developed to optimize the mean camber surface of a cranked slender wing having leading-edge vortex flow at transonic-maneuver conditions using the suction analogy. This type of flow was assumed because it was anticipated that the slenderness of the wing would preclude attached flow at the required lift coefficient. A constraint was imposed in that the camber deflections were to be restricted by a realistic structural-box requirement. The resulting application yielded mean-camber shapes which produced effective suction levels equivalent to 77 percent of the full-planar leading-edge value at the design lift coefficient.

Lamar, J. E.↗
NASA NTRS1982

A vortex wake capturing method for potential flow calculations

A method is presented for modifying finite difference solutions of the potential equation to include the calculation of non-planar vortex wake features. The approach is an adaptation of Baker's 'cloud in cell' algorithm developed for the stream function-vorticity equations. The vortex wake is tracked in a Lagrangian frame of reference as a group of discrete vortex filaments. These are distributed to the Eulerian mesh system on which the velocity is calculated by a finite difference solution of the potential equation. An artificial viscosity introduced by the finite difference equations removes the singular nature of the vortex filaments. Computed examples are given for the two-dimensional time dependent roll-up of vortex wakes generated by wings with different spanwise loading distributions.

Murman, E. M.↗
NASA NTRS2002

Rotor Wake Vortex Definition: Initial Evaluation of 3-C PIV Results of the Hart-II Study

An initial evaluation is made of extensive three-component (3C) particle image velocimetry (PIV) measurements within the wake across a rotor disk plane. The model is a 40 percent scale BO-105 helicopter main rotor in forward flight simulation. This study is part of the HART II test program conducted in the German-Dutch Wind Tunnel (DNW). Included are wake vortex field measurements over the advancing and retreating sides of the rotor operating at a typical descent landing condition important for impulsive blade-vortex interaction (BVI) noise. Also included are advancing side results for rotor angle variations from climb to steep descent. Using detailed PIV vector maps of the vortex fields, methods of extracting key vortex parameters are examined and a new method was developed and evaluated. An objective processing method, involving a center-of-vorticity criterion and a vorticity 'disk' integration, was used to determine vortex core size, strength, core velocity distribution characteristics, and unsteadiness. These parameters are mapped over the rotor disk and offer unique physical insight for these parameters of importance for rotor noise and vibration prediction.

Burley, Casey L.↗
NASA NTRS1998

Automatic Vortex Core Detection

An eigenvector method for vortex identification has been applied to recent numerical and experimental studies in external flow aerodynamics. This paper shows that it is an effective way to extract and visualize features such as vortex cores, spiral vortex breakdowns, and vortex bursts. The algorithm has also been incorporated in a finite element flow solver to guide an automatic mesh refinement program. Results show that this approach can resolve small scale vortical structures in helicopter rotor simulations which are not captured on coarse meshes.

Kenwright, David↗
NASA NTRS1982

Vortex simulation of an inviscid shear layer

The accuracy of the vortex-blob method was tested by simulating a free-shear-layer instability, Kirchhoff's elliptical vortex, and a circular vortex. The main numerical parameters in the vortex-blob method are the density of the vortices, and the distribution of vorticity within each vortex core. The growth rate of a periodic unstable mode of the shear layer was calculated numerically and compared with the exact result. The error is only a few percent for about 10 rows of vortex blobs. The error is reduced by decreasing the spacing between vortices and, correspondingly, the core size. In the simulation of the motion of the elliptical vortex, the rotation of the boundary, without change of shape, and the circular particle paths of the vortical fluid were well simulated. For the circular vortex, optimum sets of parameters were obtained by comparing them with the exact velocity. The results are consistent with convergence theories of the vortex-blob method. In particular, second-order convergence is observed with a Gaussian core from velocity calculation.

Nakamura, Y.↗
NASA NTRS1988

Modeling effects on the prediction of aerodynamic performance of a double-delta wing

A numerical investigation is undertaken of an 80/65-deg double-delta flat wing by means of the free vortex sheet method that views the vortex system on the leading edge in three different ways, in order to assess the effect of the model's type on the aerodynamic performance of the wing. The model employing two separate vortex systems, one each on inward and outward leading edges, yields predictions that are in close agreement with experimental data, although the model with single-vortex representation is adequate and requires less computational time.

Reddy, C. S.↗
NASA NTRS1976

The inviscid pressure field on the tip of a semi-infinite wing and its application to the formation of a tip vortex

A method was developed for determining the aerodynamic loads on the tip of an infinitely thin, swept, cambered semi-infinite wing at an angle of attack which is operating subsonically in an inviscid medium and is subjected to a sinusoidal gust. Under the assumption of linearized aerodynamics, the loads on the tip are obtained by superposition of the steady aerodynamic results for angle of attack and camber, and the unsteady results for the response to the sinusoidal gust. The near field disturbance pressures in the fluid surrounding the tip are obtained by assuming a dipole representation for the loading on the tip and calculating the pressures accordingly. The near field pressures are used to drive a reduced form of the Navier-Stokes equations which yield the tip vortex formation. The combined viscid-inviscid analysis is applied to determining the pressures and examining the vortex rollup in the vicinity of an unswept, uncambered wing moving steadily at a Mach number of 0.2 at an angle of attack of 0.1 rad. The viscous tip flow calculation shows features expected in the tip flow such as the qualitatively proper development of boundary layers on both the upper and lower airfoil surfaces. In addition, application of the viscous solution leads to the generation of a circular type flow pattern above the airfoil suction surface.

Hall, G. F.↗
NASA NTRS1994

Measurement Of Crossflow Vortex Structure

Method developed for measuring wavelengths of crossflow vortices by using surface-mounted, microthin, multielement hot-film sensors. Provides direct and true value of wavelength of crossflow vortices at various spanwise locations without localized flow disturbances. Attainment of laminar airflow on aircraft wings has significant potential for reducing drag and increasing fuel efficiency.

Maddalon, Dal V.↗
NASA NTRS1978

Subsonic vortex-flow design study for slender wings

A theoretical study describing the effects of spanwise camber on the lift dependent drag of slender delta wings having leading-edge-vortex-flow is presented. The earlier work by Barsby, using conical flow, indicated that drag levels similar to those in attached flow could be obtained. This is reexamined and then extended to the more practical case of nonconical flow by application of the vortex-lattice method coupled with the suction-analogy and the recently developed Boeing free-vortex-sheet method. Lastly, a design code is introduced which employs the suction analogy in an attempt to define 'optimum' camber surfaces for minimum lift dependent drag for vortex flow conditions

Lamar, J. E.↗
NASA NTRS1986

Tactical missile aerodynamics

The present conference on tactical missile aerodynamics discusses autopilot-related aerodynamic design considerations, flow visualization methods' role in the study of high angle-of-attack aerodynamics, low aspect ratio wing behavior at high angle-of-attack, supersonic airbreathing propulsion system inlet design, missile bodies with noncircular cross section and bank-to-turn maneuvering capabilities, 'waverider' supersonic cruise missile concepts and design methods, asymmetric vortex sheding phenomena from bodies-of-revolution, and swept shock wave/boundary layer interaction phenomena. Also discussed are the assessment of aerodynamic drag in tactical missiles, the analysis of supersonic missile aerodynamic heating, the 'equivalent angle-of-attack' concept for engineering analysis, the vortex cloud model for body vortex shedding and tracking, paneling methods with vorticity effects and corrections for nonlinear compressibility, the application of supersonic full potential method to missile bodies, Euler space marching methods for missiles, three-dimensional missile boundary layers, and an analysis of exhaust plumes and their interaction with missile airframes.

Hemsch, Michael J.↗
NASA NTRS1982

A vortex-filament and core model for wings with edge vortex separation

A vortex filament-vortex core method for predicting aerodynamic characteristics of slender wings with edge vortex separation was developed. Semi-empirical but simple methods were used to determine the initial positions of the free sheet and vortex core. Comparison with available data indicates that: (1) the present method is generally accurate in predicting the lift and induced drag coefficients but the predicted pitching moment is too positive; (2) the spanwise lifting pressure distributions estimated by the one vortex core solution of the present method are significantly better than the results of Mehrotra's method relative to the pressure peak values for the flat delta; (3) the two vortex core system applied to the double delta and strake wings produce overall aerodynamic characteristics which have good agreement with data except for the pitching moment; and (4) the computer time for the present method is about two thirds of that of Mehrotra's method.

Pao, J. L.↗
NASA NTRS1984

Wind tunnel wall interference in closed, ventilated and adaptive test sections

A wall interference correction method for closed rectangular test sections was developed which uses measured wall pressures. Measurements with circular discs for blockage and a rectangular wing as a lift generator in a square closed test section validate this method. These measurements are intended to be a basis of comparison for measurements in the same tunnel using ventilated (in these case, slotted) walls. Using the vortex lattice method and homogeneous boundary conditions, calculations were performed which show sufficiently high pressure levels at the walls for correction purposes in test sections with porous walls. In Gottingen, an adaptive test section (which is a deformable rubber tube of 800 mm diameter) was built and a computer program was developed which is able to find the necessary wall adaptation for interference-free measurements in a single step. To check the program prior to the first run, the vortex lattice method was used to calculate wall pressure distributions in the nonadapted test section as input data for the one-step method. Comparison of the pressure distribution in the adapted test section with free-flight data shows nearly perfect agreement. An extension of the computer program can be made to evaluate the remaining interference corrections.

Holst, H.↗
NASA NTRS1974

An Analytical Investigation of Wing-Jet Interaction

The aerodynamic interaction between the wing and an inviscid jet with Mach number nonuniformity is formulated by using a two-vortex-sheet model for the jet. One of the vortex sheets accounts for the induced jet flow and the other the induced outer flow. No additional source distribution is needed for the jet at an angle of attack. The above problem is solved by satisfying the jet and wing tangency and the jet pressure-continuity conditions and using a quasi vortex lattice method for computing the induced flow field. The latter method is derived through theoretical consideration by properly accounting for singularities present in the equations and possesses the same simplicity and generality as the conventional vortex lattice method but has a better rate of numerical convergence. The resulting system of algebraic equations is solved by Purcell's vector method. The numerical formulation is first applied to the wing-slipstream interaction problem. Results for one centered-jet configuration are compared with those predicted by some existing theories.

Lan, E. C.↗
NASA NTRS1981

A vortex-filament and core model for wings with edge vortex separation

A method for predicting aerodynamic characteristics of slender wings with edge vortex separation was developed. Semiempirical but simple methods were used to determine the initial positions of the free sheet and vortex core. Comparison with available data indicates that: the present method is generally accurate in predicting the lift and induced drag coefficients but the predicted pitching moment is too positive; the spanwise lifting pressure distributions estimated by the one vortex core solution of the present method are significantly better than the results of Mehrotra's method relative to the pressure peak values for the flat delta; the two vortex core system applied to the double delta and strake wing produce overall aerodynamic characteristics which have good agreement with data except for the pitching moment; and the computer time for the present method is about two thirds of that of Mehrotra's method.

Pao, J. L.↗
NASA NTRS1990

Application of vortex embedding to aircraft flows

A combined Eulerian/Lagrangian 'vortex embedding' method was applied to two cases. The first case involved compressible flow over a realistic fighter aircraft (X-29 fighter) at low angle of attack, for both subsonic and transonic speeds, with shock. Good agreement was obtained for computed wing surface pressures when compared to flight test data. The second case involved incompressible flow over a delta wing with round leading edges at moderate angles of attack. Strong, concentrated vortices were shown to be shed from the leading edge. The vortex embedding method appears to be an efficient solution procedure for cases where well-defined vortex sheets are shed from surfaces.

Steinhoff, John↗
NASA NTRS1991

Inverse problem in incompressible, irrotational axisymmetric flow

A vortex-sheet method for solving the axisymmetric inverse problem is presented, and an iterative, interactive computer program for computing the body shape starting from an assumed shape is developed. The method eliminates the calculation of the direct problem at every iteration using the given velocity. The singular integral that arises in the problem formulation has been integrated analytically. The efficiency of the vortex-sheet method is demonstrated using three test cases, and the obtained body shapes and the corresponding surface velocity distributions are presented.

Dinavahi, Surya P. G.↗